It has become common practice to control the mechanical aerodynamic flight surfaces of an aircraft, such as a helicopter, with an electronic control system. Operator (pilot) inputs are typically provided by position or force transducers which provide electrical or optical signals, proportional to the movement of, or force on, an input stick, to the control system. The outputs of such electronic systems typically drives 4 electro-hydraulic actuators or prime movers, one for each of the 4 main axes of the helicopter, i.e., pitch (longitudinal cyclic), roll (lateral cyclic), collective, and yaw. Three of the actuators are mechanically linked through a "mixer" to the swashplate (which controls the main rotor blades) and one is linked to the tail rotor.
Typically, the design of such control systems use feedback signals, e.g., velocity (speed), acceleration, altitude, attitude, and angular rate, from various sensors to provide negative feedback for closed loop operation or to change gains or break-point frequencies to affect control system loop response. However, these control design techniques do not provide avoidance of excessive output overshoots, which can be caused by excessive operator commands. For example, along the longitudinal axis of a helicopter, if the rotor blade tip path plane (which determines vehicle pitch attitude) approaches the front or rear of the fuselage in a rapid manner, the rotor blades may overshoot and strike the fuselage causing severe damage. This problem is exacerbated by modern systems that tend to isolate the pilot input stick command from the actual system command.
In purely mechanical systems employing mechanical input levers and pedals having displacements related to the positions of the output actuators, this problem is less likely to occur because there is a direct relationship between the position of the input levers and the real flight commands and outputs. Furthermore, in mechanical systems there is built-in damping and slew rate limiting based on the time constant of the inertia of the thing being controlled, thus, adding more "feel" to the input lever. Similarly, for a system employing an electrical input command signal proportional to lever position, there is also less isolation because the pilot always knows the position of the swashplate by the position of the lever.
It is known that a side-arm controller such as that described in U.S. Pat. No. 4,420,808, allows the pilot to control all 4 axes of the aircraft with a single input stick. The controller comprises a limited travel (approx 0.25 in) control grip stick which provides inputs for all axes to the electronic control system by sensing a force applied to the stick in a given direction using strain gauge sensors. Typically, the force applied to the stick is proportional to a rate command, i.e., a rate of change in the loop reference, along a given control axis, e.g., deg/sec for vehicle pitch attitude. When no force is applied to the stick, the rate command is zero, the loop reference is fixed, and the control system holds all the vehicle axes at the current reference settings.
Because the inertia of the control grip of the side-arm controller is negligible and the input stick position relates to a rate of change command rather than, say, a position command, it is possible for the pilot to inadvertently command a high rate of change (slew rate) causing the output actuator to be driven into an undesirable operating region. This occurs because the input lever position does not relate directly to real air-vehicle output position of the swashplate and tail rotor blade pitch. Thus, the pilot can easily lose the "feel" of the actual command to, and output position of, the system (swashplate and tail rotor).
This is particularly a problem for axes which exhibit an inertial time lag between the pilot command and the aircraft response, e.g., pitch attitude, which can cause the pilot to make a larger input than is necessary.